Low pressure recovery inlet for gas turbine engine turbomachinery on aircraft
AssigneeRolls-Royce North American Technologies Inc.
InventorsAlan W. Smith, Michael C. Moser, Chadd D. Engel, Bryan H. Lerg
Life of this patent
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Abstract
A gas turbine engine system adapted for use with an aircraft includes a gas turbine engine. The gas turbine engine system includes a gas turbine engine and an auxiliary compressor operably coupled with the gas turbine engine. The auxiliary compressor arranged to receive air from an inlet that faces away from a direction of travel of the aircraft such that the air flowing to the auxiliary compressor has a lower mass flow rate of air at an inflight airspeed than a mass flow rate of air of a forward facing inlet so that the auxiliary compressor consumes less work from the turbine inflight.
Claims19
A gas turbine engine system adapted for use with an aircraft, the gas turbine engine system comprising:
- a gas turbine engine including a gearbox and an engine core having a first compressor, a combustor, and a turbine coupled with the first compressor and the gearbox to drive the first compressor and the gearbox,
- a second compressor operably coupled with the gas turbine engine and driven by the turbine via the gearbox during operation of the gas turbine engine to produce a flow of auxiliary compressed air usable by the aircraft, and
- a nacelle extending from a fore end and an aft end thereof, a direction of a flow of air over the aircraft flowing from the fore end to the aft end, the nacelle arranged around the gas turbine engine and formed to define (i) a first inlet at the fore end of the nacelle and arranged to direct a first portion of the flow of air aft into the first compressor, (ii) an outlet formed in the aft end of the nacelle and configured to direct exhaust gases from the turbine away from the nacelle, and (iii) a second inlet in direct fluid communication with ambient air and arranged to direct a second portion of the flow of air into the second compressor, the second portion of the flow of air being separate from the first portion of the flow of air,
- wherein the second inlet is oriented such that the second portion of the flow of air flows through the second inlet in a direction defined by the aft end of the nacelle toward the fore end of the nacelle.
A gas turbine engine system adapted for use with an aircraft, the gas turbine engine system comprising:
- a gas turbine engine including a gearbox and an engine core having a first compressor, a combustor, and a turbine,
- a second compressor operably coupled with the gas turbine engine and driven by the turbine via the gearbox, and
- an airframe including fuselage, a wing, and a nacelle, a nose and a tail disposed aft of the nose such that a direction of a flow of air over the airframe moves from the nose of the airframe aft towards the tail, the nacelle extending from a fore end and an aft end thereof, the nacelle arranged around the gas turbine engine and formed to define a first inlet at the fore end of the nacelle and arranged to direct a first portion of the flow of air aft into the first compressor, and an outlet formed in the aft end of the nacelle and configured to direct exhaust gases from the turbine away from the nacelle,
- wherein the airframe is formed to define a second inlet in direct fluid communication with ambient air and arranged to direct a second portion of the flow of air into the second compressor, the second portion of the flow of air being separate from the first portion of the flow of air, the second inlet is oriented such that the second portion of the flow of air flows through the second inlet in a direction defined by the aft end of the nacelle toward the fore end of the nacelle.
Dependencies are inferred from the claim text.